Modular Thermal Management System for Spacecraft

ABSTRACT

A composite panel provides structural strength and rigidity for modular assembly of spacecraft while serving the dual purposes of structure and heat transfer for thermal management of an environment for equipment, such as a spacecraft. Instruments, gimbals, surveillance, imaging, detectors, and the like may be mounted in a spacecraft designed and constructed from standard panels to provide the structural and heat transfer requirements to support the onboard equipment. Extremely small temperature differentials required by the panels support a substantially isothermal perimeter for the structure, able to sink heat from any location on a panel, transport it to a rejection site, and reject it to the cold environment of space, thus easing the task of design and packaging of instrumentation and infrastructure of satellites and other spacecraft. Two-phase mass transport of fluids inside the panels aids high heat flux rates with minimal temperature differentials, even with reinforced, composite, polymeric materials for the structural elements of the panels.

RELATED APPLICATIONS

This application: (1) is a continuation-in-part of co-pending U.S. patent application Ser. No. 11/743,555, filed on May 2, 2007; (2) claims the benefit of co-pending U.S. Provisional Patent Application Ser. No. 60/954,550, filed on Aug. 7, 2007; and (3) claims the benefit of co-pending U.S. Provisional Patent Application Ser. No. 61/042,205, filed on Apr. 3, 2008.

BACKGROUND

1. The Field of the Invention

This invention relates to heat transfer and, more particularly, to novel systems and methods for combined structural and heat transfer panels for spacecraft.

2. The Background Art

Heat transfer is fundamental to many processes. Engines releasing energy from fuel must routinely reject heat. Meanwhile, electrical and electronic devices that consume electricity likewise require cooling. One particularly demanding environment occurs with spacecraft. Spacecraft contain many electronic instruments. Weight is at a premium, as is space. Meanwhile, the loads required of any structure during launch may be substantial. Thus, the combination of light weight and high strength is difficult to achieve.

In heat exchange, metals have comparatively high thermal conductivities. By contrast, composite materials such as bonded composites, fiber-reinforced layers, and the like typically have comparatively poor thermal conductivity. Design and optimization of size, weight, strength, stiffness, location, packaging, and heat transfer systems is typically unique to every space craft. Thus lead times are long, and reuse or salvage value of any design is minimal. A modular system is needed in which standard components may be assembled to provide a stock spacecraft suitable for hosting equipment for a particular mission. Reusable components would also be an advance in the art.

What is needed is an apparatus and method for modular, lightweight, structurally strong and stiff, dual purpose panels providing both structure and heat exchange for a spacecraft without additional structures for that purpose. It would be an advance in the art to provide heat exchangers that also provide structural functionality to the apparatus. Moreover, it would be an advance in the art to provide a mechanism for maintaining as nearly as possible an isothermal belt about the circumference or other perimeter of a satellite in order to provide temperature stability and better dwell times or “time on target” for an observation platform to maintain its orientation without having to change in order to avoid solar radiation loads.

BRIEF SUMMARY OF THE INVENTION

In view of the foregoing, in accordance with the invention as embodied and broadly described herein, a method and apparatus are disclosed in one embodiment of the present invention as including a modular, dual-function, structural-thermal-integrated spacecraft system.

An integrated structural and heat transfer panel relies on the two-phase thermal panels, and variations thereof, the basic operational and configuration details thereof having already been disclosed in U.S. patent application Ser. No. 11/743,555 filed May 2, 2007, U.S. Provisional Patent Application Ser. No. 60/954,550, filed on Aug. 7, 2007; and U.S. Provisional Patent Application Ser. No. 61/042,205, filed on Apr. 3, 2008, all of which are incorporated herein by reference. Panels in accordance with the invention integrate thermal management as well as structural requirements, particularly for satellite and other spacecraft systems.

In one embodiment, the panels are manufactured in a modular configuration, rather than in specialized, unique configurations. Thus, several panels may be used to create an assembly that approximates a cube or other rectangular structure. Each panel of the system may function independently from other panels, but may yet be capable of maintaining thermal contact with other panels at a sufficiently robust rate to substantially maintain a large portion, and even the entire perimeter of the satellite at substantially a single temperature.

For example, temperature differentials ranging from tens to hundreds of degrees exist to drive heat in various heat transfer mechanisms and devices. By contrast, panels in accordance with the invention may maintain temperature differentials of as little as two degrees Fahrenheit or less between a heat sink location receiving heat from a device mounted thereto and a location of heat rejection by the panel to the environment. Thus, individual panels may be capable of maintaining a temperature throughout within such a narrow limit. Meanwhile, whether or not a particular panel is viewing the sun or black space, interconnected panels may stabilize and reduce temperature differentials while efficiently moving heat about the connected system of panels.

Moreover, individual panels may be provided with doors having a thermally reflective side and a thermally emissive side. Thus, when a door is closed over a thermal panel, the exposed area reflects thermal radiation received from the sun or other objects. Meanwhile, when the door opens, the emissive side of the door operates like one of the fixed panels, while the now exposed, fixed panel also operates as an emissive panel.

Flexible thermal links, pivoting thermal links, or the like, thermally connecting doors to fixed panels forming the periphery of the satellite or other spacecraft may provide a comparatively high rate of heat transfer therebetween, thus providing significant radiation augmentation by the door panels.

In one embodiment, each panel has at least one face or surface that has a high emissivity finish so as to function effectively as a space radiator. Panels may be connected at their edges, or may themselves be formed to turn a corner and provide an additional segment, something like a foot extending from the main leg of the principal portion of the panel. Thus, the foot of a panel may provide, and be designed to provide, a sufficient area to provide substantial heat transfer at minimal temperature differential between the working fluid within one panel and the working fluid of another adjacent panel attached thereto.

In certain embodiments, panels need not be thermally connected to one another. Thus, each individual panel may operate independently. In yet other embodiments, the doors of a panel may provide a reflective surface, a solar electric device, or both when exposed to a radiation load, such as the sun. The doors may be opened to provide an emissive surface when exposed to the cold blackness space (e.g., when shaded). Thus, thermal loads being absorbed by panels in accordance with the invention may be transferred to another place on the same panel or to another connected panel to be rejected according to the orientation and exposure of the satellite or other spacecraft.

Doors may be configured in various sizes. For example, a pair of doors may have a central opening about half way across a surface of a fixed panel being covered. Thus, the two doors together may effectively cover a fixed panel of a satellite. When the two doors are opened fully to be substantially flat and coplaner with respect to the previously covered fixed panel, the effective radiation space or surface area may effectively be double that of the fixed panel alone. By contrast, when the doors are closed against the fixed panel, the missive surfaces of the door panels are basically re-reflecting back into the fixed panel, while the outer surfaces of the doors are reflecting solar radiation, or collecting radiation if a solar electric device is present. Thus, the effective radiating ability may be controlled, and may be multiplied or diminished, even to approach negligible values. When exposed to solar radiation, a fixed panel may be covered by the door panels presenting a reflective surface, solar electric cell array, or both. When shaded or facing black space, a fixed panel may be exposed by an open door for radiating heat away from the satellite, and may be augmented by additional heat transfer radiating from the emissive surfaces of any door.

In another embodiment, multiple doors may cover a fixed panel, each covering it entirely. For example, from one side of the panel, a door may swing toward a fixed panel to cover it. Meanwhile, from the opposite edge of the fixed panel, another door may close across substantially the entire radiative area of the same fixed panel. Thus, an outer door covers an inner door, which in turn directly covers the fixed panel. Thus, the exposed, radiating, surface area exposed to a cold or shaded space environment might be about triple the area of a single fixed panel alone. Four such doors may conceivably be folded over a rectangular, fixed panel to provide about five times the radiating area of the panel alone.

Nevertheless, in certain embodiments, a satellite or other spacecraft may simply contain, for example, six sides of a cube, none shielded from incoming radiation and all having the same emissive properties and substantially the same areas. In such an embodiment, heat loads from the sun may be driven by very low temperature differentials from the irradiated surfaces of the spacecraft to the shaded sides thereof to be re-radiated to the coldness of space.

In certain embodiments of an apparatus and method in accordance with the invention, a satellite or other spacecraft may be designed and built from modular panel components. For example, standardized panels may form the sides of a rectangular box or a cube. Ends may be substantially similar, with relief or other finishing treatments applied to fit and close corners and other irregular areas. Thus, the box formed by a standardized panel structure may be assembled from several basic panels that all may have substantially the same configuration. The net inventory of distinct part numbers may be substantially reduced.

Moreover, such a “box” may be fitted with equipment internally positioned against the panels without regard to the precise thermal loads of the contained equipment. For example, because the heat transfer and the minimal temperature differentials required for heat transfer are so effective, any rack or piece of electrical or mechanical equipment may be properly connected mechanically and thermally to a wall panel of the box. Heat will be properly rejected from the outside surface of the panels after transfer from the heat source (equipment) to a location exposed to the cold environment of space for proper rejection. No need remains for complex paths, careful positioning, complex analysis of relative locations, and additional weight for conductive thermal paths internal to the spacecraft.

Some of the features provided by such an apparatus include efficient and effective transport of heat across, around, and even through a spacecraft. In certain embodiments, the panels may effectively create isothermal zones within the spacecraft when desirable and so designed. In some embodiments, the entire perimeter of a spacecraft or a single panel of the spacecraft may be rendered nearly isothermal within a matter of a couple of degrees or less. Meanwhile, the radiating surfaces of the panels may be particularly efficient having substantially uniform temperatures thereacross.

Due to the structure of the panels, mass is greatly reduced in a spacecraft by the dual functioning of the structural panels as heat transfer panels, and vice versa. Thus, networks of thermal passages, radiation auxiliarly panels, and the like are not needed. The need for careful allocation of space may be greatly relieved by rendering as much of the satellite as desired, including substantially the entire outer perimeter as a radiating surface if desired.

In certain embodiments, substantially the entire structure of a spacecraft, satellite, or the like may be integrated with the thermal panel in the same physical structure. Accordingly, all operating equipment may be attached to the internal surface of each panel in the structure to sink heat into the combination structural, thermal panels. Meanwhile, the outer surfaces of those same thermal panels may then reject heat to the comparatively space environment.

In certain embodiments, the integration of structural and thermal functions in a single physical element can greatly reduce the mass required for heat rejection panels, solar electrical collection panels, heat transfer paths internal to a satellite, structural framing, and so forth. Specifically, when compared to solid metal conduction or radiation surfaces, a light weight “sandwich” panel in accordance with the invention may require substantially less metal, and may even be constructed from a composite material such as a carbon-fiber-reinforced, resin matrix, operating as both an effective heat transfer mechanism and as a stiff and strong structural member.

In accordance with certain aspects of an apparatus and method in accordance with the invention, mechanical and thermal stability result from greater stiffness and lower coefficients of thermal expansion available in composite materials. By contrast, solid metals, with their isotropic thermal and mechanical properties cannot achieve the same directional properties available by proper design with a composite material.

Reduction of the mass of radiating surfaces, in some cases elimination of all extraneous masses outside of the structures, facilitates both the deployment and positioning of radiating surfaces with minimum impact on spacecraft stability. Likewise, certain configurations of the radiating surfaces may be adjusted from nearly zero to several times the surface area of any particular facet of a spacecraft. For example, doors may be connected to a fixed panel by a flexible or pivoting thermal link to provide minimal temperature differentials thereacross. Meanwhile, up to four rectangular panels of substantially the same size as one fixed panel may conceal it by folding over one another when not in use. They may then be opened to present a total of about five times the effective area of the fixed panel available to radiate.

With so much control over the surface area exposure of a fixed panel, a spacecraft module may effectively be switched on or off to thermal radiation, reflection, or both. Thus, a panel may be placed in an energy dissipation mode, an energy absorbing mode, or a neutral adiabatic mode.

Because operational components that typically produce heat may be mounted directly on an interior surface of a panel in accordance with the invention, spreading heat to radiating surfaces becomes substantially automatic. Meanwhile, to the extent that a spacecraft does not need all surfaces for heat transfer, solar power arrays may occupy part of the outer surfaces thereof. Moreover, by the use of doors, solar power arrays may be disposed on exterior surfaces thereof, to operate simultaneously on opposite faces thereof and with radiating fixed panels. Likewise, solar power arrays may be placed on reflective doors in order to collect electrical power while denying thermal inputs into fixed panels having doors closed thereover.

Thus, in certain embodiments, an apparatus and method in accordance with the invention may combine and optimize the simultaneous deployment of solar power arrays, heat radiating functions, and may include shielding, insulation, reflection, and other functionalities in any combination or proportionality desired. At a minimum, the available surface area available for heat rejection from a spacecraft may be enormously greater than the mere perimeter surfaces of the spacecraft at launch.

Stabilization of the thermal profile of a spacecraft to substantially isothermal conditions throughout the perimeter reduces the temperature extremes to which individual internal components may be exposed. In certain embodiments, the heat sink temperature available from a panel sinking heat from a heat generating component may be sufficiently stable that such a device may operate full time, independent of orientation or exposure to the sun. For example, the need to shut devices down because their solar radiation has increased too much may be greatly mediated or eliminated by simply transferring the heat away therefrom around the substantially isothermal belt of panels of the satellite or other spacecraft. Because the apparatus may actually go into an adiabatic mode, on a component basis, or on an overall spacecraft basis, survival heaters may be reduced or completely eliminated. That is, although solar heat load is often a problem for certain detector devices such as infrared focal-plane arrays, some components may suffer from too much cooling. An apparatus and method in accordance with the invention may provide modulated cooling and heating by engineered operation of reflective (on face) and emissive (opposite face) doors over emissive, fixed panels. Such designs may benefit from a standard configuration of a basic panel, readily adapted to assembly as a standardized box. The box may satisfy both mechanical and thermal requirements to support a variety of internal payloads. It may greatly reduce costs of design, fabrication, inventory maintenance, repairs, and so forth.

In certain embodiments of apparatus in accordance with the invention, multiple spacecraft may be assembled as generic boxes. Although a cube is one embodiment, other embodiments are contemplated having arbitrary lengths with modular panels assembled to interact for mechanical and thermal functionality.

Detectors, such as surveillance instruments may have greater “time on target” because they need not be so sensitive to solar radiation. Whether shielded by selectively openable and closeable doors, or simply relying on a substantially isothermal belt carrying heat from a heated region to a cold, heat-rejecting region of the paneled box, potential heat loads may be rejected from the satellite or other spacecraft by radiating from fixed, door, or both panels viewing space from the shadowed or shaded side of a spacecraft. Addition of reflective doors or surfaces may simply augment this reduction of solar load and reduce the requirement for heat rejection required for a spacecraft.

BRIEF DESCRIPTION OF THE DRAWINGS

The foregoing features of the present invention will become more fully apparent from the following description and appended claims, taken in conjunction with the accompanying drawings. Understanding that these drawings depict only typical embodiments of the invention and are, therefore, not to be considered limiting of its scope, the invention will be described with additional specificity and detail through use of the accompanying drawings in which:

FIG. 1 is a perspective view of one embodiment of a spacecraft constructed from dual-service, structural-thermal-integrated panels in accordance with the invention;

FIG. 2 is a perspective view of one embodiment of a spacecraft in accordance with the invention having one end panel removed, and provided with an aperture in a side panel for supporting a surveillance or other detector device in the spacecraft;

FIG. 3 is an orthogonal view of one embodiment of an end panel for an apparatus such as those illustrated in FIGS. 1 and 2;

FIG. 4 is a cross-sectional view of the end panel of FIG. 3;

FIG. 5 is an exploded view of one embodiment of an apparatus in accordance with the invention, relying on multiple faces formed by a single continuous panel having bends, to thus avoid and reduce the need for additional conductive paths and connections between panels;

FIG. 6 is a top plan view of one embodiment of a spacecraft in accordance with the invention having doors selectively openable to selectively shield radiation panels or to augment radiation panels, including reliance on flexible thermal links to carry heat from fixed panels to door panels;

FIG. 7 is a top plan view of the embodiment of FIG. 5 showing the door panels in an alternative configuration of deployment with reflective panel faces toward solar incidence and radiative, heat rejecting faces of panels toward a cold space environment;

FIGS. 8 and 9 are orthogonal views of connected panels showing flexible thermal links installed;

FIG. 10 is a top plan view of one embodiment of a fixed panel linked to be protected by a reflective surface of a door panel, alternatively closed to shield or opened to expose radiating fixed panels, the door panels being connected to the fixed panel by a flexible thermal link and a deployment connection and driving mechanism;

FIG. 11 is an end view of the apparatus of FIG. 10 in an partially open, deployed configuration;

FIG. 12 is a top plan view of an alternative embodiment of a spacecraft having door panels that each cover substantially one entire fixed panel, and including an internal panel suitable for mounting hardware that may generate heat, and transferring that heat through a foot on an opposing end of an internal “cross panel” outward toward an external fixed panel of the spacecraft; and

FIG. 13 is a top plan view of one embodiment of the apparatus of FIG. 12 illustrating some alternative locations of heat producing components positioned both internally with respect to the spacecraft, as well as externally secured to door panels in accordance with the invention.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

It will be readily understood that the components of the present invention, as generally described and illustrated in the drawings herein, could be arranged and designed in a wide variety of different configurations. Thus, the following more detailed description of the embodiments of the system and method of the present invention, as represented in the drawings, is not intended to limit the scope of the invention, as claimed, but is merely representative of various embodiments of the invention. The illustrated embodiments of the invention will be best understood by reference to the drawings, wherein like parts are designated by like numerals throughout.

Referring to FIG. 1, and to FIGS. 1-13 generally, a system 10 in accordance with the invention may include various panels 12 providing dual-service, structural-thermal, integrated, modular panels 12. These panels are disclosed in detail in various selected embodiments in the documents, incorporated herein by reference.

In the illustrated embodiment, side panels 12 a, 12 b, 12 c, 12 d are all substantially identical. That is, each may be manufactured in a modular fashion to be connected by a thermal link or may include both a leg portion 14 and a foot portion 16. The foot portion 16 and leg portion of each panel 12 are contiguous and continuous both mechanically and thermally, including the internal fluid communication of the operating fluid. Thus, the leg portion 14 is substantially isothermal with respect to the foot portion 16 (within the operating temperature differential thereacross).

Suitable bonding methods, including welding, bolting, thermal grease contact, framing, and other fastening mechanisms. Such may be relied on in according with good engineering design to provide intimate thermal contact between the toe 16 of any particular panel 12, and an adjacent leg portion 14 of another panel 12 in thermal contact therewith. Thus, by means of a single, highly efficient, contact portion through the foot 16 of a panel 12, each panel 12 may transfer heat to the next adjacent panels 12 in contact with it.

Mechanically, the structural strength and stiffness of the system 10 is assured by the material and structural integrity of the leg portion 14 and foot portion 16 of each panel 12. Thus, the interlocking or fitted panels 12, each may support tensile, compressive, and sheer forces within the basic plane of the panel 12, as well as within the foot 16. Thus, suitable structural stiffness and strength may continue throughout all of the panels 12 in the apparatus 10.

In the illustrated embodiment, the top panel 12 e is matched by a bottom panel 12, not shown. Because of the shape of modular side panels 12 such as the panel 12 a, the top panel 12 e may be relieved to fit the characteristic shape of the remainder of the apparatus 10. Accordingly, the top panel 12 e and the bottom panel 12 may each have one or more foot portions 16 in contact with, for example opposing panels 12 a, 12 c, or opposing panels 12 b, 12 d.

Referring to FIG. 2, in one embodiment of an apparatus in accordance with the invention, racks may be, configured either open or enclosed as shown, and may be mounted directly to panels 12 in order to obtain heat-sink services from the panels 12. Meanwhile, the racks 18 receive structural service in terms of strength, stiffness, rigidity, and other mechanical functions by virtue of the structural properties of the panels 12.

The foot portion 16 of each of the panels 12 may be sized to overlap a sufficient distance to provide structural stiffness, shear support, and sufficient thermal contact area to maintain a substantially isothermal belt, within a matter of a few degrees or less, about the perimeter of the apparatus 10. In the illustrated embodiment, the racks 18 a, 18 b are mounted against opposing side panels 12 d, 12 b, respectively. Nevertheless racks 18 may be mounted to any particular panel 12, including side panels 12 a, 12 c, and top panels 12 e or bottom panels 12 e.

In the embodiment of FIG. 2, an aperture 20 provides support for a lens, camera, focal-plane array, or other surveillance or tracking device associated with the rack 18 a. Thus, for example, a focal-plane array detector mounted in the rack 18 a may have visibility through an aperture 20, and have structural, mounting services and thermal transfer services provided by the side panel 12 b into the system 10. Each of the panels 12 operates in accordance with the heat and mass transport principles of the panel design 12 with its internal working fluid operating in two phases in accordance with the principles of heat pipes, and typically wickless liquid transport along the interior surfaces of the panel 12.

FIG. 3 illustrates an end panel 12 e suitable for the top or bottom of the system 10. In the illustrated embodiment, a leg portion 14 may have one, two, or more foot portions 16 in contact with the side panels 12. Relief 22 provided in the end panel 12 e may be designed to fit the foot portions 16 of side panels 12.

With discontinuities at the corners of the end panel 12 e, a panel 12 e may be fabricated to have a foot portion 16 along each edge as continuous and continuous parts of the panel 12 e. Thus, each of the feet 16 may maintain fluid communication and be unitary with the leg portion 14. Feet 16 may be sized to optimize the heat transfer, weight, structural stiffness and strength, as well as temperature differentials in order to maintain the overall perimeter of the apparatus 10 in a substantially isothermal condition.

By isothermal is meant that an apparatus has but a single temperature. By substantially isothermal is meant that within the range of the temperature differential acceptable to the operation of the apparatus 10, the temperature variations about the perimeter of the apparatus 10 remains substantially within or effectively within that range.

For example, in the apparatus 10 illustrated, an entire exposed surface of a panel 12 may effectively represent a single temperature within a matter of a few degrees Fahrenheit or less. Meanwhile, approximately the same temperature differential may be required to transfer heat between a foot portion 16 of one panel 12, and the adjacent foot portion 16 of an adjacent panel 12.

Nevertheless, material thicknesses, material compositions, wall thicknesses, and the like may be designed in order to optimize temperature gradients. For example, in one embodiment, walls of foot portions 16 may be thinner than the main leg portion 14 at corresponding points. Meanwhile, an interfacing segment of a leg portion 14 in contact with a foot portion 16 may likewise be thinned sufficiently to render the conductive path substantially equivalent to a single wall thickness of the leg portion 14 of the rest of a panel 12.

Since a fluid heat transfer coefficient is acting on either side of the conductive path at such a joint, temperature differentials between adjacent panels 12 may typically not be identical to the temperature differential in the fluid across the expanse of a single panel 12. Nevertheless, engineering design can provide temperature profiles suitable for meeting the demands of onboard equipment racks 18, and may be designed to provide a “substantially isothermal” environment. Such a system 10 may have temperature differentials on the order of a few degrees, as opposed to tens or hundreds of degrees of difference between locations of heat reception by panels 12 and heat rejection into the surrounding space environment by panels 12.

Referring to FIG. 4, in one embodiment of an apparatus 10 in accordance with the invention, the foot portions 16 may simply be extended to become other panels 12 f, 12 g. A frame 28 may optionally be provided, or the corresponding edge 24, 26 of adjacent panels 12 f, 12 g may be welded, or fastened by some other suitable means. Thermal contact may be improved between adjacent panels 12 f, 12 g, by various mechanisms. Nevertheless, in the illustrated embodiment, the apparatus 10 may be formed as a substantially rigid rectangular volume. Possible shapes include a cube, or a rectangular box of some other aspect ratio between the areas of the various faces thereof. In the illustrated embodiment, the frame 28 may simply be a solid of suitable material and configuration, or may be formed of heat pipes or panels 12 sized appropriately, just as the foot portions 16 may be formed. Thus, the frame 28, itself, in certain embodiments may actually be a two-phase fluid-containing, heat-transfer device. In other embodiments, the frame 28 may itself be a box having solid sides.

Likewise, panels may contain channels as grids, previously disclosed. Alternatively, panels may be segmented with separating walls between ranks of channels. Thus, a rack may connect across two or more segmented ranges of a panel. Thus, penetration of a single portion of a panel 12 by a high-speed particle does not disable heat transfer from a rack 18 into the panel 12.

Referring to FIGS. 5-6, in one embodiment of an apparatus 10 in accordance with the invention, panels 12 may be provided with doors 30. Doors may be connected to adjacent fixed panels 12 by flexible thermal links 32, pivoting thermal links 32, or the like. In general, any time flexible thermal links 32 are mentioned, pivoting thermal links 32 may be considered as an alternative included therein. These links 32 may be applied to any of the doors 30 illustrated herein. The number and size thereof may be engineered to provide a desired heat flux. Flexible thermal links may provide thermal contact, even intimate thermal contact, while providing mechanical isolation. Thus, a flexible thermal link 32 may permit a door 30 to pivot with respect to a fixed panel 12, while still maintaining substantially intimate thermal contact therebetween.

In the illustrated embodiment, the doors 30 may each include a highly emissive surface 34 on an interior face thereof, and a highly reflective surface 36 on an outer face thereof. Accordingly, when a door 30 is closed over a fixed panel 12, the apparatus 10 presents a highly reflective surface 36 to the environment. Also, a solar cell panel may be positioned on the reflective face to provide power from solar radiation.

By contrast, when one or more doors 30 of the apparatus 10 are opened, they may present a highly emissive surface 34 to the surrounding space environment. In the embodiment of FIG. 5, two fixed panels 12 are enclosed, or have their doors 30 closed on them. Accordingly, those surfaces 36, if presented to the sun, reflect solar radiation. Meanwhile, adjacent doors 30, when deployed, also reflect any solar radiation coming from the same direction.

Meanwhile, the opposite half of the apparatus 10 having doors 30, open provides radiative surfaces toward a shaded (e.g., cold sky or space) environment. Both the fixed panels 12, and their adjacent, open doors radiate to the environment. Thus fixed panels 12, and the highly emissive surfaces 34 of the deployed panels 30 radiate to the cold or shaded environment of sky or space. Meanwhile, the reflective surfaces 36 along with those of the closed doors 30 may be designed to reflect incoming solar radiation coming from the opposite direction.

Referring to FIG. 6, the apparatus 10 may have or present a solar radiation environment with three or more sides of the apparatus 10 having high 14 reflectivity surfaces. Meanwhile, by deploying fully open doors 30, one side thereof provides the high reflectivity surface 36 to the solar radiation environment. Meanwhile, the opposite face presents a high emissity surface 34, along with that of the fixed panel 12, to the shaded or cold sky of a space environment.

Referring to FIG. 7, in one embodiment of an apparatus 10 in accordance with the invention, a fixed panel 12 may have various, even infinite, positioning for the doors 30. For example, in the embodiment of FIG. 7, the doors 30 are closed on the fixed panel 12. Meanwhile, the fixed panel 12 is thermally connected and mechanically isolated with respect to the doors 30 by the flexible thermal links 32. A suitable deployment mechanism 38 may provide a pivoting relationship, a driver, and typically both to position the doors 30 with respect to the fixed panel 12.

Referring to FIGS. 8-9, the panel 12 of FIG. 7 with the doors 30 in a partially or fully open position provides a reflective surface 36, while the front of the door 30 provides a highly emissive surface 34. The thermal links 32 maintain thermal contact to render the emissive surface 34 almost equivalent to that of the fixed panel 12 as explained hereinabove.

The doors 30 may be deployed to be substantially coplanar with the fixed panel 12. According to the particular needs of each particular face of an apparatus 10 in accordance with the invention, the doors 30 may be designed and operated to control solar influx of heat, and thermal radiation out to cold space or a shaded environment.

Referring to FIGS. 10-11, in one embodiment of an apparatus in accordance with the invention, a panel 12, or a fixed panel 12 may be connected by flexible thermal links 38 to multiple doors 30, each covering substantially the entire area or extent of the fixed panel 12. Accordingly, each of the two doors 30 may cover either the fixed panel 12, directly, or the other door 30. Upon deployment, one door 30 may be opened, maintaining the other door 30 closed over the fixed panel 12. Thus, a single panel face 34 may radiate, while all other exposed surfaces are reflecting. Meanwhile, upon opening of the second door 30, then more and more of the fixed panel 12 as well as the high emissivity surface 34 of the door 30 may be exposed. Ultimately, the configuration of FIGS. 10-11 may provide almost triple the heat rejection capacity that a single fixed panel 12 may provide. Similarly, four doors 30 may be deployed about a single panel 12.

Referring to FIGS. 12-13, in one embodiment of an apparatus 10 in accordance with the invention, a system 10 may include several fixed panels 12 forming the structure thereof. In the illustrated embodiment, doors 30 hinge from corners of the apparatus 10, and may include deployment mechanisms and flexible thermal links 32 as appropriate. Meanwhile, certain racks 18 may actually be attached to doors 30, to be hidden within the apparatus 10 when the door 30 is closed, and exposed when the door 30 is opened. Meanwhile, other racks 18 may be placed inside the apparatus 10 and connected to the interior surfaces of the fixed panels 12.

In the illustrated embodiment, an additional internal panel 12 (e.g., cross panel) may extend between opposite fixed panels 12 of the apparatus 10. Accordingly, additional internal racks 18 connected to the cross-connected, internal panel 12 may reject heat into that internal panel 12, to be rejected or transported through foot portions 16 thereof into the fixed panels 12 connected at either end thereof.

In the illustrated embodiment, substantially the entire perimeter of the apparatus 10 may be radiating, reflecting, or various faces or surfaces thereof may be selectively opened or closed to radiate or reflect, respectively according to whether or not that particular face of the apparatus 10 is called upon to reflect away solar radiation, or reject heat to the shaded environment of cold space.

The present invention may be embodied in other specific forms without departing from its spirit or essential characteristics. The described embodiments are to be considered in all respects only as illustrative, and not restrictive. The scope of the invention is, therefore, indicated by the appended claims, rather than by the foregoing description. All changes which come within the meaning and range of equivalency of the claims are to be embraced within their scope. 

1. A spacecraft comprising: a first panel forming a first container comprising opposing layers, formed of a structural material and provided with channels transporting a liquid phase of an operating fluid by capillary action toward a comparatively hotter location within the first container and vapor spaces transporting a vapor phase of the operating fluid by convection from the comparatively hotter location toward a comparatively cooler location within the first container; a second panel forming a second container comprising opposing layers, formed of a structural material and provided with channels transporting a liquid phase of an operating fluid by capillary action toward a comparatively hotter location within the second container and vapor spaces transporting a vapor phase of the operating fluid by convection from the comparatively hotter location toward a comparatively cooler location within the second container; the second panel connected substantially rigidly to the first panel and in thermal communication therewith to exchange heat therewith and forming a structural unit therewith; and a mount connecting a source of heat to at least one of the first and second panels structurally and thermally to exchange mechanical force and heat therewith.
 2. The spacecraft of claim 1, further comprising third and fourth panels, the first, second, third, and fourth panels connected substantially rigidly to the first panel and in thermal communication therewith to exchange heat therewith and forming a structural unit therewith.
 3. The spacecraft of claim 2, wherein the third panel forms a third container comprising opposing layers, formed of a structural material and provided with channels transporting a liquid phase of an operating fluid by capillary action toward a comparatively hotter location within the third container and vapor spaces transporting a vapor phase of the operating fluid by convection from the comparatively hotter location toward a comparatively cooler location within the third container.
 4. The spacecraft of claim 3, wherein the forth panel forms a third container comprising opposing layers, formed of a structural material and provided with channels transporting a liquid phase of an operating fluid by capillary action toward a comparatively hotter location within the forth container and vapor spaces transporting a vapor phase of the operating fluid by convection from the comparatively hotter location toward a comparatively cooler location within the forth container.
 5. The spacecraft of claim 4, wherein the first, second, third, and fourth panels form the primary frame structure of the spacecraft to which other components of the spacecraft connect.
 6. The spacecraft of claim 5, wherein the first, second, third, and fourth panels are substantially isothermal.
 7. The spacecraft of claim 6, wherein the first, second, third, and fourth panels form a perimeter of the spacecraft. 